CN118128603B - Turbine disk system cooling sealing structure - Google Patents
Turbine disk system cooling sealing structure Download PDFInfo
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- CN118128603B CN118128603B CN202410572149.0A CN202410572149A CN118128603B CN 118128603 B CN118128603 B CN 118128603B CN 202410572149 A CN202410572149 A CN 202410572149A CN 118128603 B CN118128603 B CN 118128603B
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- 238000001816 cooling Methods 0.000 title claims abstract description 87
- 238000007789 sealing Methods 0.000 title claims abstract description 39
- 244000126211 Hericium coralloides Species 0.000 claims description 22
- 238000000034 method Methods 0.000 claims description 7
- 210000004210 tooth component Anatomy 0.000 claims description 4
- 230000000694 effects Effects 0.000 abstract description 12
- 230000001939 inductive effect Effects 0.000 abstract description 3
- 238000009423 ventilation Methods 0.000 description 4
- 230000009286 beneficial effect Effects 0.000 description 2
- 238000004378 air conditioning Methods 0.000 description 1
- 238000010586 diagram Methods 0.000 description 1
- 230000003116 impacting effect Effects 0.000 description 1
- 230000005476 size effect Effects 0.000 description 1
- 238000006467 substitution reaction Methods 0.000 description 1
Classifications
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
- F01D5/02—Blade-carrying members, e.g. rotors
- F01D5/08—Heating, heat-insulating or cooling means
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D11/00—Preventing or minimising internal leakage of working-fluid, e.g. between stages
- F01D11/02—Preventing or minimising internal leakage of working-fluid, e.g. between stages by non-contact sealings, e.g. of labyrinth type
- F01D11/04—Preventing or minimising internal leakage of working-fluid, e.g. between stages by non-contact sealings, e.g. of labyrinth type using sealing fluid, e.g. steam
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- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
Abstract
The application provides a turbine disk system cooling and sealing structure, which belongs to the technical field of aeroengines and comprises a first-stage turbine disk, a first-stage turbine lower comb plate, a first-stage turbine rear upper sealing comb assembly and a first-stage turbine rear comb assembly, wherein the first-stage turbine lower comb plate, the first-stage turbine rear upper sealing comb assembly and the first-stage turbine rear comb assembly are connected with the first-stage turbine lower comb plate, and one end of the first-stage turbine rear comb assembly, which is far away from the first-stage turbine disk, is connected with a second-stage turbine disk; the device also comprises a first-stage turbine guide vane right support ring, a first-stage turbine guide vane left support ring and a guide disc which are connected end to end, wherein the lower end of the guide disc is also connected with a honeycomb ring which is in clearance fit with a first-stage turbine lower comb plate; the left support ring of the first-stage turbine guide vane is provided with a plurality of air-inducing holes, one side of the right support ring of the first-stage turbine guide vane, which is close to the first-stage turbine disc, is provided with an impact cooling hole, and the position of the first-stage turbine disc, which corresponds to the direction of the lower air flow, is provided with an air vent. According to the scheme, the cooling sealing effect and the air entraining quality of the air system are improved, the size of the short-service-life engine is reduced, and the performance of the engine is improved.
Description
Technical Field
The application relates to the technical field of aeroengines, in particular to a turbine disc system cooling sealing structure.
Background
The disposable short-life engine has the characteristics of small size, low cost, high rotating speed and small air flow, and the smaller size and air flow cause 'size effect' in the aspects of aerodynamics, structure, strength and the like, so that the difficulty is increased for the aerodynamic design under the small size, and meanwhile, the challenges are brought to the rotor-stator cooling design. Therefore, the problem of thermal cooling caused by the rotation effect is more remarkable, and the air flow of the engine is small, so that the cooling air-entraining amount of the air system of the engine is more required to be strictly controlled, and the performance of the engine is ensured under the condition of less air-entraining as much as possible. The air quantity of the air system directly influences the cooling and sealing effect of the rotor and the stator, and further influences the strength life of the engine, so that various measures are required to be taken to improve the cooling and sealing quality of the air system.
The conventional turbine disk sealing cooling flow path is divided into a plurality of parts after the air at the outlet of a proper amount of air compressor is accelerated and depressurized through a pre-rotation nozzle, so that the turbine disk is sealed and cooled, and the edge of the turbine disk is sealed; for blades with cooling designs, the turbine blades also need to be cooled. The short-life engine usually adopts a design without cooling blades, and is limited by size and cost, and a complicated pre-rotation cooling system is not adopted, so that a high-efficiency and simple turbine disk system cooling sealing structure and method are required to be developed to meet specific use requirements of the short-life engine.
Disclosure of Invention
In view of the above, the embodiment of the application provides a turbine disk system cooling and sealing structure, which can adapt to the characteristics of small size, small flow, short service life, low cost and the like of a short-service-life engine under the condition of small bleed air flow, realize efficient cooling and sealing of the turbine disk system of the short-service-life engine, simplify the structural design of the engine, improve the bleed air sealing and cooling quality of an air system, and reduce the development cost of the engine.
The embodiment of the application provides a turbine disk system cooling and sealing structure, which comprises a first-stage turbine disk, wherein the front end of the first-stage turbine disk is connected with a first-stage turbine lower comb plate, the rear end of the first-stage turbine disk is connected with a first-stage turbine rear upper sealing comb assembly and a first-stage turbine rear comb assembly, and one end of the first-stage turbine rear comb assembly, which is far away from the first-stage turbine disk, is connected with a second-stage turbine disk;
The device also comprises a first-stage turbine guide vane right support ring, a first-stage turbine guide vane left support ring and a guide disc, wherein the first-stage turbine guide vane right support ring, the first-stage turbine guide vane left support ring and the guide disc are used for forming a cavity body in an end-to-end connection mode, the first-stage turbine guide vane left support ring is fixedly connected with the high-pressure turbine guide vane, the lower end of the guide disc is also connected with a honeycomb ring, and the honeycomb ring is in clearance fit with a lower grate disc of the first-stage turbine; the left support ring of the first-stage turbine guide vane is provided with a plurality of air-inducing holes, one side of the right support ring of the first-stage turbine guide vane, which is close to the first-stage turbine disk, is provided with an impact cooling hole, a front upper airflow cavity of the first-stage turbine disk is formed between the right support ring of the first-stage turbine guide vane and the first-stage turbine disk, a lower airflow cavity of the first-stage turbine disk is formed between the guide disk and the honeycomb ring, cooling airflow passing through the impact cooling hole is divided into upper airflow flowing through the front upper airflow cavity of the first-stage turbine disk and lower airflow flowing through the lower airflow cavity of the first-stage turbine disk, vent holes are formed in positions, corresponding to the directions of the lower airflow, of the first-stage turbine disk, and cooling airflow passing through the vent holes passes through the first-stage turbine disk, and then the upper sealing comb assembly and the first-stage turbine rear comb assembly flow out respectively.
According to a specific implementation mode of the embodiment of the application, the rear upper sealing comb tooth assembly of the first-stage turbine disk comprises a single comb tooth and a first-stage turbine disk on a second-stage turbine guide vane supporting ring, the single comb tooth on the second-stage turbine guide vane supporting ring is arranged at the upper end of the second-stage turbine guide vane supporting ring close to the first-stage turbine disk, the second-stage turbine guide vane supporting ring is arranged on the second-stage turbine guide vane, the single comb tooth on the second-stage turbine guide vane supporting ring and the right end of the first-stage turbine disk form clearance fit, and the other end of the second-stage turbine guide vane supporting ring and the rear comb tooth assembly of the first-stage turbine are clearance fit.
According to a specific implementation of the embodiment of the application, an included angle is arranged between the axis of the bleed hole and the vertical line of the left support ring of the first-stage turbine guide vane.
According to a specific implementation of the embodiment of the application, the included angle between the axis of the bleed hole and the perpendicular to the left support ring of the primary turbine vane is in the range of 20 ° -30 °.
According to a specific implementation of an embodiment of the application, the bleed holes have a diameter ofThe number of the air holes isThe area of the air-inducing hole is A 1, and the diameter of the impact cooling hole isThe number of the impact cooling holes isThe area of the impingement cooling holes is A 2, and the diameter of the ventilation holes isThe number of the vent holes isWherein, the method comprises the steps of, wherein,，，，。
According to a specific implementation mode of the embodiment of the application, the gap between the honeycomb ring and the lower grate plate of the primary turbine ranges from 0.07mm to 0.13mm.
According to a specific implementation of an embodiment of the application, the clearance between the second stage turbine vane support ring and the first stage turbine rear grate assembly ranges from 0.06mm to 0.14mm.
According to a specific implementation manner of the embodiment of the application, the hole edge of the vent hole is provided with a rounding structure.
The beneficial effects are that:
According to the turbine disk cooling sealing structure, a turbine guide vane right support ring, a guide vane left support ring, a guide disk and a first-stage turbine disk form a first-stage turbine disk front sealing flow path, and a first-stage turbine disk, a second-stage turbine guide vane support ring, a single comb tooth, a first-stage turbine rear comb tooth assembly and a second-stage turbine guide vane support ring form a first-stage turbine disk rear sealing cooling flow path; the primary turbine disk is subjected to impact cooling through the arrangement of the bleed holes, so that heat exchange is enhanced, and the cooling effect is improved under the condition of low bleed air flow; and the air holes are directly opened on the turbine disk, so that the air flow in front of the first-stage turbine disk can be led to the first-stage turbine disk, the air-entraining path can be effectively shortened, the flow path structure is simplified, the along-path loss and the air-entraining temperature rise are reduced, and the sealing and cooling effects behind the turbine disk are improved. Therefore, the structure can greatly improve the cooling and sealing effects of the air system, improve the air entraining quality, and is beneficial to reducing the size of the short-service-life engine and improving the performance of the engine.
Drawings
In order to more clearly illustrate the technical solutions of the embodiments of the present application, the drawings that are needed in the embodiments will be briefly described below, and it is obvious that the drawings in the following description are only some embodiments of the present application, and that other drawings can be obtained according to these drawings without inventive effort for a person skilled in the art.
FIG. 1 is a schematic view of a turbine disk cooling seal structure according to an embodiment of the present invention;
FIG. 2 is a schematic diagram of a cooling airflow path of a turbine disk cooling seal structure according to an embodiment of the invention;
FIG. 3 is a layout of impingement cooling holes of a turbine disk cooling seal structure in accordance with an embodiment of the present invention.
In the figure: 1. a first stage turbine vane right support ring; 2. a first stage turbine vane left support ring; 3. an air vent; 4. a deflector disc; 5. a honeycomb ring; 6. a first-stage turbine lower comb plate; 7. a primary turbine disk; 8. a vent hole; 9. impingement cooling holes; 10. a second stage turbine vane support ring; 11. single comb teeth on the second-stage turbine guide vane supporting ring; 12. a first stage turbine rear grate assembly; 13. an upper airflow cavity in front of the first-stage turbine disk; 14. a primary turbine disk lower airflow cavity; 15. the rear upper airflow cavity of the first-stage turbine disk; 16. a secondary turbine disk; 17. and the rear upper part of the primary turbine disc is tightly sealed with the comb tooth component.
Detailed Description
Embodiments of the present application will be described in detail below with reference to the accompanying drawings.
Other advantages and effects of the present application will become apparent to those skilled in the art from the following disclosure, which describes the embodiments of the present application with reference to specific examples. It will be apparent that the described embodiments are only some, but not all, embodiments of the application. The application may be practiced or carried out in other embodiments that depart from the specific details, and the details of the present description may be modified or varied from the spirit and scope of the present application. It should be noted that the following embodiments and features in the embodiments may be combined with each other without conflict. All other embodiments, which can be made by those skilled in the art based on the embodiments of the application without making any inventive effort, are intended to be within the scope of the application.
It is noted that various aspects of the embodiments are described below within the scope of the following claims. It should be apparent that the aspects described herein may be embodied in a wide variety of forms and that any specific structure and/or function described herein is merely illustrative. Based on the present disclosure, one skilled in the art will appreciate that one aspect described herein may be implemented independently of any other aspect, and that two or more of these aspects may be combined in various ways. For example, an apparatus may be implemented and/or a method practiced using any number of the aspects set forth herein. In addition, such apparatus may be implemented and/or such methods practiced using other structure and/or functionality in addition to one or more of the aspects set forth herein.
It should also be noted that the illustrations provided in the following embodiments merely illustrate the basic concept of the present application by way of illustration, and only the components related to the present application are shown in the drawings and are not drawn according to the number, shape and size of the components in actual implementation, and the form, number and proportion of the components in actual implementation may be arbitrarily changed, and the layout of the components may be more complicated.
In addition, in the following description, specific details are provided in order to provide a thorough understanding of the examples. However, it will be understood by those skilled in the art that the aspects may be practiced without these specific details.
The embodiment of the application provides a cooling and sealing structure of a turbine disk system, which can realize efficient cooling and sealing of the turbine disk system under the condition of small bleed air flow, and can simplify the structure of an engine and reduce the cost of the engine while meeting the use requirement, and is described in detail below with reference to fig. 1 to 3.
Referring to fig. 1 and 2, the cooling seal structure of the turbine disk system of the present embodiment includes a primary turbine disk 7, a primary turbine lower grate disk 6 is connected to the front end of the primary turbine disk 7, a primary turbine disk rear upper seal grate assembly 17 and a primary turbine rear grate assembly 12 are connected to the rear end of the primary turbine disk 7, and one end of the primary turbine rear grate assembly 12 away from the primary turbine disk 7 is connected to a secondary turbine disk 16;
The device further comprises a first-stage turbine guide vane right support ring 1, a first-stage turbine guide vane left support ring 2 and a guide disc 4 which are used for forming a cavity body and connected end to end, wherein the first-stage turbine guide vane left support ring 2 is fixedly connected with the high-pressure turbine guide vane, the lower end of the guide disc 4 is also connected with a honeycomb ring 5, and the honeycomb ring 5 is in clearance fit with a first-stage turbine lower comb plate 6; the left support ring 2 of the first-stage turbine guide vane is provided with a plurality of air introducing holes 3, one side of the right support ring 1 of the first-stage turbine guide vane, which is close to the first-stage turbine disk 7, is provided with an impact cooling hole 9, a front upper air flow cavity 13 of the first-stage turbine disk is formed between the right support ring 1 of the first-stage turbine guide vane and the first-stage turbine disk 7, a lower air flow cavity 14 of the first-stage turbine disk is formed between the guide disk 4 and the honeycomb ring 5, cooling air flow passing through the impact cooling hole 9 is divided into upper air flow flowing through the front upper air flow cavity 13 of the first-stage turbine disk and lower air flow flowing through the lower air flow cavity 14 of the first-stage turbine disk, an air vent 8 is arranged at a position of the first-stage turbine disk 7 corresponding to the direction of the lower air flow, and the cooling air flow passing through the air vent 8 passes through the upper sealing comb assembly 17 and the rear comb assembly 12 of the first-stage turbine disk respectively flows out.
In the embodiment, a front sealing flow path of the first-stage turbine disk 7 is formed by the first-stage turbine guide vane right support ring 1, the first-stage turbine guide vane left support ring 2, the guide disk 4 and the first-stage turbine disk 7; the primary turbine disk 7, the single-grate teeth 11 on the secondary turbine guide vane support ring, the primary turbine rear-grate tooth component 12 and the secondary turbine guide vane support ring 10 form a rear-seal cooling flow path of the primary turbine disk 7. For the first-stage turbine guide vane right support ring 1, the first-stage turbine guide vane left support ring 2 and the flow guiding disk 4 which form the cavity head-to-tail connection, all the components are in sealing connection, and the main connection structure is as follows: one end of the first-stage turbine guide vane left support ring 2 is connected with the first-stage turbine guide vane right support ring 1, the other end of the first-stage turbine guide vane left support ring 2 is connected with a guide disc 4, the upper end of the guide disc 4 is connected with the first-stage turbine guide vane right support ring 1, and the lower end of the guide disc 4 is connected with a honeycomb ring 5, so that cooling air flows entering from the air entraining holes 3 flow out through the impact cooling holes 9. Through setting up bleed hole 3 to carrying out the impingement cooling to one-level turbine dish 7, strengthen the heat transfer, promote cooling effect under the little bleed flow condition.
Secondly, be equipped with the interval between one-level turbine stator right branch collar 1 and the one-level turbine dish 7, form the air conditioning flow path, consequently, the cooling air current that impingement cooling hole 9 flowed out divide into the upper air current that flows through the preceding upper air current chamber 13 of one-level turbine dish and the lower air current that flows through the lower air current chamber 14 of one-level turbine dish, after setting up air vent 8, after the cooling air current that impingement cooling hole 9 flowed out will be partly discharged to one-level turbine dish 7 through air vent 8, consequently, directly open air vent 8 on the turbine dish, can effectively shorten the bleed air route after leading the air current in front of one-level turbine dish 7 to one-level turbine dish 7, simplify the flow path structure, reduce along journey loss and bleed air temperature rise, promote and the sealed cooling effect behind the turbine dish.
In addition, the lower grate disk 6 of the primary turbine is fixed on the primary turbine disk 7, the left end of the rear grate component 12 of the primary turbine is fixed on the primary turbine disk 7, and the right end is connected with the secondary turbine disk 16. Through the arrangement method, the cooling flow paths can be tightly sealed before and after the primary turbine disk 7, and the leakage flow of air flow can be effectively controlled by controlling the clearance of the comb tooth assembly.
In one embodiment, the upper rear sealing comb assembly 17 of the primary turbine disk comprises a secondary turbine guide vane support ring 10 and a single comb tooth 11 on the secondary turbine guide vane support ring, the single comb tooth 11 on the secondary turbine guide vane support ring is arranged at the upper end of the secondary turbine guide vane support ring close to the primary turbine disk 7, the secondary turbine guide vane support ring 10 is arranged on the secondary turbine guide vane, the single comb tooth 11 on the secondary turbine guide vane support ring forms clearance fit with the right end of the primary turbine disk 7, and the other end of the secondary turbine guide vane support ring 10 is in clearance fit with the primary turbine rear comb tooth assembly 12. The rear end of the primary turbine disc 7 and the outer part of the secondary turbine guide vane support ring 10 form a primary turbine disc rear upper airflow cavity 15, and cooling airflow flowing out of the vent holes 8 is discharged from the primary turbine disc rear upper airflow cavity 15 through a gap between the single grate tooth 11 on the secondary turbine guide vane support ring and the primary turbine disc 7. The single comb tooth 11 on the second-stage turbine guide vane support ring is positioned on the stator, namely the single comb tooth 11 on the second-stage turbine guide vane support ring and the second-stage turbine guide vane support ring 10 are of an integrated structure, when the engine actually runs, the deformation of the rotor stator is mutually coordinated, the axial movement of the comb tooth sealing assembly can be effectively prevented, and the sealing effect is ensured.
In one embodiment, an included angle is formed between the axis of the gas introduction hole 3 and the perpendicular to the left support ring 2 of the primary turbine vane. The included angle between the axis of the air introducing hole 3 and the vertical line of the left support ring 2 of the first-stage turbine guide vane is 20-30 degrees. In the embodiment, the bleed holes 3 with included angles can reduce bleed loss and improve bleed quality.
In one embodiment, the diameter of the vent hole 3 isThe number of the air entraining holes 3 isThe area of the bleed holes 3 is A 1,The impingement cooling holes 9 have a diameter ofThe number of the impingement cooling holes 9 isThe area of the impingement cooling holes 9 is a 2,The diameter of the vent hole 8 isThe number of the vent holes 8 isThe area of the vent hole 8 is A 3,Wherein, the method comprises the steps of, wherein,，，，. The entire flow path in the present embodiment forms impingement cooling, the impingement cooling holes 9 are uniformly arranged in the circumferential direction of the engine, referring to fig. 3, and the holes are small in number, form impingement cooling to the primary turbine disk 7, and enhance the cooling effect.
In one embodiment, the clearance between the honeycomb ring 5 and the primary turbine lower grate disk 6 ranges from 0.07mm to 0.13mm, and the clearance between the secondary turbine vane support ring 10 and the primary turbine rear grate assembly 12 ranges from 0.06mm to 0.14mm. If the flow rate of the bleed air is lower than the lower limit, the bleed air flow rate cannot meet the cooling and sealing requirements, and exceeds the upper limit, so that the bleed air flow rate is larger, and the performance of the engine is not facilitated.
According to a specific implementation of the embodiment of the application, the hole edge of the vent hole 8 is provided with a rounded structure. The vent hole 8 can be directly formed in the first-stage turbine disk 7, and the hole edge is of a round structure, so that the air flow before the first-stage turbine disk 7 is directly communicated to the first-stage turbine disk 7, the ventilation path is shortened, the ventilation structure is simplified, the air flow loss along the path is reduced, the air flow edge Cheng Wensheng is reduced, and the sealing and cooling after the first-stage turbine disk 7 are more facilitated.
For the cooling airflow path of the turbine disk system cooling seal structure of the present application, referring to fig. 2, the cooling airflow path includes the combustor outer ring bleed air passing through the bleed air holes 3, after impacting the cooling holes 9, the primary turbine disk 7 is impact cooled, and then the airflow is divided into three: one of the two air flows upwards through the front upper airflow cavity 13 of the first-stage turbine disk and then seals the front edge of the first-stage turbine disk 7; the second strand is discharged from the gap between the honeycomb ring 5 and the first-stage turbine lower comb plate 6 after going down to the first-stage turbine lower air flow cavity 14; the third air flow is split into two parts rightward through the vent holes 8 on the first-stage turbine disk 7, upwards passes through the upper air flow cavity 15 behind the first-stage turbine disk to seal the rear rim of the first-stage turbine disk 7 and then is converged into the main flow passage, and downwards passes through the gap between the second-stage turbine guide vane support ring 10 and the first-stage turbine rear grate assembly 12 and is discharged into the front end of the second-stage turbine.
According to the embodiment provided by the invention, the engine structure and the ventilation flow path can be simplified through the cooling structure, the air entraining quality of the engine is improved, the cooling effect is enhanced, the engine is facilitated to be simplified in structure, the size is reduced, and the engine performance is improved.
The foregoing is merely illustrative of the present application, and the present application is not limited thereto, and any changes or substitutions easily contemplated by those skilled in the art within the scope of the present application should be included in the present application. Therefore, the protection scope of the application is subject to the protection scope of the claims.
Claims (8)
1. The turbine disk system cooling and sealing structure is characterized by comprising a first-stage turbine disk (7), wherein the front end of the first-stage turbine disk (7) is connected with a first-stage turbine lower comb plate (6), the rear end of the first-stage turbine disk (7) is connected with a first-stage turbine disk rear upper sealing comb assembly (17) and a first-stage turbine rear comb assembly (12), and one end, far away from the first-stage turbine disk (7), of the first-stage turbine rear comb assembly (12) is connected with a second-stage turbine disk (16);
The device further comprises a first-stage turbine guide vane right support ring (1), a first-stage turbine guide vane left support ring (2) and a guide disc (4) which are used for forming the end-to-end connection of the cavity, wherein the first-stage turbine guide vane left support ring (2) is fixedly connected with the high-pressure turbine guide vane, the lower end of the guide disc (4) is also connected with a honeycomb ring (5), and the honeycomb ring (5) is in clearance fit with a first-stage turbine lower comb disc (6); a plurality of air introducing holes (3) are formed in the left support ring (2) of the first-stage turbine guide vane, an impact cooling hole (9) is formed in one side, close to the first-stage turbine disc (7), of the right support ring (1) of the first-stage turbine guide vane, an upper air flow cavity (13) of the first-stage turbine disc is formed between the right support ring (1) of the first-stage turbine guide vane and the first-stage turbine disc (7), a lower air flow cavity (14) of the first-stage turbine disc is formed between the guide disc (4) and the honeycomb ring (5), cooling air flow passing through the impact cooling hole (9) is divided into upper air flow flowing through the upper air flow cavity (13) of the first-stage turbine disc and lower air flow flowing through the lower air flow cavity (14) of the first-stage turbine disc, an air vent (8) is formed in a position, corresponding to the direction of the lower air flow, and cooling air flow passing through the air vent (8) passes through the first-stage turbine disc and then passes through the upper sealing comb tooth component (17) of the first-stage turbine disc and then the comb tooth component (12) respectively flows out.
2. The turbine disk cooling and sealing structure according to claim 1, wherein the upper sealing comb assembly (17) behind the first-stage turbine disk comprises a second-stage turbine guide vane support ring (10) and a single comb tooth (11) on the second-stage turbine guide vane support ring, the single comb tooth (11) on the second-stage turbine guide vane support ring is arranged at the upper end of the second-stage turbine guide vane support ring (10) close to the first-stage turbine disk (7), the second-stage turbine guide vane support ring (10) is arranged on the second-stage turbine guide vane, the single comb tooth (11) on the second-stage turbine guide vane support ring at the upper end of the second-stage turbine guide vane support ring (10) forms clearance fit with the right end of the first-stage turbine disk, and the other end of the second-stage turbine guide vane support ring (10) is in clearance fit with the first-stage turbine rear comb tooth assembly (12).
3. The turbine disk cooling seal structure according to claim 1, characterized in that an angle is provided between the axis of the bleed holes (3) and the perpendicular to the left support ring (2) of the primary turbine vane.
4. A turbine disc system cooling seal according to claim 3, characterized in that the angle between the axis of the bleed holes (3) and the perpendicular to the left support ring (2) of the primary turbine vane is in the range of 20 ° -30 °.
5. Turbine disk cooling-seal structure according to claim 1, characterized in that the bleed holes (3) have a diameter ofThe number of the air entraining holes (3) isThe area of the air entraining holes (3) is A 1, and the diameter of the impact cooling holes (9) isThe number of the impingement cooling holes (9) isThe area of the impingement cooling hole (9) is A 2, and the diameter of the vent hole (8) isThe number of the vent holes (8) isWherein, the method comprises the steps of, wherein,，，，。
6. The turbine disk cooling seal structure according to claim 1, wherein a gap between the honeycomb ring (5) and the primary turbine lower labyrinth plate (6) ranges from 0.07mm to 0.13mm.
7. The turbine disk system cooling seal structure of claim 1, wherein a clearance between the secondary turbine vane support ring (10) and the primary turbine rear grate assembly (12) ranges from 0.06mm to 0.14mm.
8. The turbine disk cooling sealing structure according to claim 1, wherein the hole edge of the vent hole (8) is provided with a rounded structure.
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| CN202410572149.0A CN118128603B (en) | 2024-05-10 | 2024-05-10 | Turbine disk system cooling sealing structure |
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| CN120007444B (en) * | 2025-03-26 | 2025-11-21 | 中国航发沈阳发动机研究所 | A bearing sealing and cooling structure for turbine casing of an aero-engine |
| CN120367663B (en) * | 2025-06-25 | 2025-08-19 | 西北工业大学 | A honeycomb bionic structure turbine disk for an aircraft engine |
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| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| GB1525746A (en) * | 1974-11-27 | 1978-09-20 | Gen Electric | Gas turbine engines including labyrinth seals |
| CN115244271A (en) * | 2020-03-24 | 2022-10-25 | 赛峰飞机发动机公司 | Turbofan engine including means for regulating the flow rate of cooling fluid |
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| Publication number | Priority date | Publication date | Assignee | Title |
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| US7743613B2 (en) * | 2006-11-10 | 2010-06-29 | General Electric Company | Compound turbine cooled engine |
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| CN116537895B (en) * | 2023-07-04 | 2023-09-15 | 中国航发四川燃气涡轮研究院 | Pre-rotation air supply system with comb gap control |
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2024
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| GB1525746A (en) * | 1974-11-27 | 1978-09-20 | Gen Electric | Gas turbine engines including labyrinth seals |
| CN115244271A (en) * | 2020-03-24 | 2022-10-25 | 赛峰飞机发动机公司 | Turbofan engine including means for regulating the flow rate of cooling fluid |
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